Blade/disk dovetail backcut for blade/disk stress reduction (7FA, stage 2)

ABSTRACT

A turbine blade that may include an airfoil and a blade dovetail, the blade dovetail being shaped corresponding to a dovetail slot in a turbine disk. The blade dovetail may include a dovetail backcut sized and positioned according to blade geometry to maximize a balance between stress reduction on the disk, stress reduction on the blade, a useful life of the turbine blade, and maintaining or improving the aeromechanical behavior of the turbine blade. A start point of the dovetail backcut, which defines a length of the dovetail backcut along a dovetail axis, may be determined relative to a datum line positioned a fixed distance from a forward face of the blade dovetail along a centerline of the dovetail axis. The start point of the dovetail backcut may be at least approximately 1.215 inches in an aft direction from the datum line.

BACKGROUND OF THE INVENTION

The present invention relates to gas turbine technology and, moreparticularly, to a modified blade and/or disk dovetail designed todivert the blade load path around a stress concentrating feature in thedisk on which the blade is mounted and/or a stress concentrating featurein the blade itself.

Certain gas turbine disks include a plurality of circumferentiallyspaced dovetails about the outer periphery of the disk defining dovetailslots therebetween. Each of the dovetail slots receives in an axialdirection a blade formed with an airfoil portion and a blade dovetailhaving a shape complementary to the dovetail slots.

The blades may be cooled by air entering through a cooling slot in thedisk and through grooves or slots formed in the dovetail portions of theblades. Typically, the cooling slot extends circumferentially 360°through the alternating dovetails and dovetail slots.

It has been found that interface locations between the blade dovetailsand the dovetail slots are potentially life-limiting locations due tooverhanging blade loads and stress concentrating geometry. In the past,dovetail backcuts have been used in certain turbine engines to relievestresses. These backcuts, however, were minor in nature and wereunrelated to the problem addressed here. Moreover, the locations andremoved material amounts were not optimized to maximize a balancebetween stress reduction on the disk, stress reduction on the blades,and a useful life of the blades.

BRIEF DESCRIPTION OF THE INVENTION

The present application thus describes a method for reducing stress onat least one of a turbine disk and a turbine blade, wherein a pluralityof turbine blades are attachable to the disk, and wherein each of theturbine blades includes a blade dovetail engageable in acorrespondingly-shaped dovetail slot in the disk. The method mayinclude: (a) determining a start point for a dovetail backcut relativeto a datum line, the start point defining a length of the dovetailbackcut along a dovetail axis; (b) determining a cut angle for thedovetail backcut; and (c) removing material from at least one of theblade dovetail or the disk dovetail slot according to the start pointand the cut angle to form the dovetail backcut. The start point and thecut angle may be optimized according to blade and disk geometry tomaximize a balance between stress reduction on the disk, stressreduction on the blade, a useful life of the turbine blades, andmaintaining or improving the aeromechanical behavior of the turbineblade. The datum line may be positioned a fixed distance from a forwardface of the blade dovetail along a centerline of the dovetail axis, andstep (a) may be practiced such that the start point of the dovetailbackcut is at least approximately 1.215 inches in an aft direction fromthe datum line.

In some embodiments, each of the turbine blades may be configured tooperate within a second stage of a 7FA turbine. Step (b) may bepracticed such that the cut angle is a maximum of about 3°. In otherembodiments, step (b) may be practiced such that the cut angle is about2°. The fixed distance from the forward face of the blade dovetail forthe datum line may be approximately 2.817 inches.

In some embodiments, optimizing of the start point and the cut angle maybe practiced by executing finite element analyses on the blade and diskgeometry. Step (b) may be practiced by determining multiple cut anglesto define the dovetail backcut with a non-planar surface. Step (c) maybe practiced by removing material from the blade dovetail. In otherembodiments, step (c) may be practiced by removing material from thedisk dovetail slot. Step (c) also may be practiced by removing materialfrom the blade dovetail and from the disk dovetail slot. Step (c)further may practiced such that a resulting angle based on the materialremoved from the blade dovetail and the disk dovetail slot does notexceed the cut angle.

The present application further describes a turbine blade that mayinclude an airfoil and a blade dovetail, the blade dovetail being shapedcorresponding to a dovetail slot in a turbine disk. The blade dovetailmay include a dovetail backcut sized and positioned according to bladegeometry to maximize a balance between stress reduction on the disk,stress reduction on the blade, a useful life of the turbine blade, andmaintaining or improving the aeromechanical behavior of the turbineblade. A start point of the dovetail backcut, which defines a length ofthe dovetail backcut along a dovetail axis, may be determined relativeto a datum line positioned a fixed distance from a forward face of theblade dovetail along a centerline of the dovetail axis. The start pointof the dovetail backcut may be at least approximately 1.215. inches inan aft direction from the datum line.

In some embodiments, each of the turbine blades may be configured tooperate within a second stage of a 7FA turbine. A cut angle of thedovetail backcut may be a maximum of about 3°. In certain embodiments, acut angle of the dovetail backcut may be about 2°. The fixed distancefrom the forward face of the blade dovetail for the datum line may beapproximately 2.817 inches. In some embodiments, the dovetail backcutmay have a non-planar surface.

The present application further describes a turbine blade that includesan airfoil and a blade dovetail, the blade dovetail being shapedcorresponding to a dovetail slot in a turbine disk. Each of the turbineblades may be configured to operate within a second stage of a 7FAturbine and the blade dovetail includes a dovetail backcut. A startpoint of the dovetail backcut, which defines a length of the dovetailbackcut along a dovetail axis, may be determined relative to a datumline positioned a fixed distance from a forward face of the bladedovetail along a centerline of the dovetail axis. The start point of thedovetail backcut may be at least approximately 1.215 inches in an aftdirection from the datum line. A cut angle of the dovetail backcut maybe a maximum of about 3°.

In some embodiments, a cut angle of the dovetail backcut may be about2°. The fixed distance from the forward face of the blade dovetail forthe datum line may be approximately 2.817 inches. These and otherfeatures of the present application will become apparent upon review ofthe following detailed description of the preferred embodiments whentaken in conjunction with the drawings and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of an exemplary gas turbine disk segmentwith attached gas turbine blade;

FIG. 2 is a perspective view of the pressure side of the exemplary gasturbine blade;

FIG. 3 is a perspective view of the suction side of the exemplary gasturbine blade;

FIGS. 4-7 illustrate close-up views of blade or disk dovetail areas inwhich material will be removed;

FIGS. 8 and 9 illustrate a material removal area for a stage 1 blade ordisk in a first turbine class of a first type (a 6FA turbine);

FIGS. 10 and 11 illustrate a material removal area for a stage 1 bladeor disk in the first turbine class of a second type (a 6FA+e turbine);

FIG. 12 shows a material removal area for a stage 2 blade or disk in thefirst turbine class of the second type (the 6FA+e turbine);

FIGS. 13 and 14 illustrate a material removal area for a stage 1 bladeor disk in a second turbine class of a first type (a 7FA+e turbine);

FIG. 15 shows a material removal area for a pressure side of a stage 2blade or disk in the second turbine class of the first type (the 7FA+eturbine);

FIG. 16 shows a material removal area for a suction side of the stage 2blade or disk in the second turbine class of the first type (the 7FA+eturbine);

FIGS. 17 and 18 illustrate a material removal area for a stage 1 bladeor disk in a third turbine class of a first type (a 9FA+e turbine);

FIG. 19 shows a material removal area for a pressure side of a stage 2blade or disk in the third turbine class of the first type (the 9FA+eturbine);

FIG. 20 shows a material removal area for a suction side of the stage 2blade or disk in the third turbine class of the first type (the 9FA+eturbine);

FIGS. 21-30 illustrate the determination of datum line W for each stageblade or disk of each turbine class and type.

FIGS. 31 and 32 illustrate a material removal area for a stage 1 bladeor disk in the third turbine class of a second type (a 9FA turbine);

FIGS. 33 and 34 illustrate a material removal area for a stage 1 bladeor disk in the second turbine class of a second type (a 7FA turbine);and

FIGS. 35 and 36 illustrate a material removal area for a stage 2 bladeor disk in the second turbine class of the second type (the 7FAturbine).

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a perspective view of an exemplary gas turbine disk segment 10in which is secured a gas turbine blade 12. The gas turbine disk 10includes a dovetail slot 14 that receives a correspondingly shaped bladedovetail 16 to secure the gas turbine blade 12 to the disk 10. FIGS. 2and 3 show opposite sides of a bottom section of the gas turbine blade12 including an airfoil 18 and the blade dovetail 16. FIG. 2 illustratesa so-called pressure side of the gas turbine blade 12, and FIG. 3illustrates a so-called suction side of the gas turbine blade 12.

The dovetail slots 14 are typically termed “axial entry” slots in thatthe dovetails 16 of the blades 12 are inserted into the dovetail slots14 in a generally axial direction, i.e., generally parallel but skewedto the axis of the disk 10.

An example of a gas turbine disk stress concentrating feature is thecooling slot. The upstream or downstream face of the blade and disk 10may be provided with an annular cooling slot that extendscircumferentially a full 360°, passing through the radially innerportion of each dovetail 16 and dovetail slot 14. It will be appreciatedthat when the blades are installed on the rotor disk 10, cooling air(e.g., compressor discharge air) is supplied to the cooling slot whichin turn supplies cooling air into the radially inner portions of thedovetail slots 14 for transmittal through grooves or slots (not shown)opening through the base portions of the blades 12 for cooling theinterior of the blade airfoil portions 18.

A second example of a gas turbine disk stress concentrating feature isthe blade retention wire slot. The upstream or downstream face of theblade 12 and disk 10 may be provided with an annular retention slot thatextends circumferentially a full 360°, passing through the radiallyinner portion of each dovetail 16 and dovetail slot 14. It will beappreciated that when the blades are installed on the rotor disk 10, ablade retention wire is inserted into the retention wire slot which inturn provides axial retention for the blades.

The features described herein are generally applicable to any airfoiland disk interface. The structure depicted in FIGS. 1-3 is merelyrepresentative of many different disk and blade designs across differentclasses of turbines. For example, at least three classes of gas turbinesincluding disks and blades of different sizes and configurations aremanufactured by General Electric Company (“GE”) of Schenectady, N.Y.These include: (1) a first turbine class, GE's 6FA and 6FA+e turbines;(2) a second turbine class, GE's 7FA and 7FA+e turbines; and (3) a thirdturbine class, GE's 9FA and 9FA+e turbines. Each turbine additionallyincludes multiple stages within the turbine having varying blade anddisk geometries.

It has been discovered that the interface surfaces between the bladedovetail 16 and the disk dovetail slot 14 are subject to stressconcentrations that are potentially life-limiting locations of theturbine disk 10 and/or turbine blade 12. It would be desirable to reducesuch stress concentrations to maximize the life span of the disk and/orblade without negatively impacting the life span or aeromechanicalbehavior of the gas turbine blades.

With reference to FIGS. 4-7, the gas turbine blade dovetail 16 includesa number of pressure faces or tangs 20 on the dovetail pressure side anda number of pressure faces or tangs 20 on the dovetail suction side.Depending on the turbine class and blade and disk stage, a backcut 22may be made on either or both of the suction side aft end and pressureside forward end of the blade dovetail tangs 20 or disk dovetail tangs21 (see FIG. 1). With particular reference to FIGS. 6 and 7, the backcut22 is formed by removing material from the pressure faces 20 of theblade dovetail 16 or disk dovetail slot 14. The material can be removedusing any suitable process such as a grinding or milling process or thelike, which may be the same as or similar to the corresponding processused for forming the blade dovetail 16 or disk dovetail slot 14.

The amount of material to be removed and thus the size of the backcut 22is determined by first determining a start point for the dovetailbackcut relative to a datum line, the start point defining the length ofthe dovetail backcut along the dovetail axis. A cut angle is alsodetermined for the dovetail backcut, the exemplary angle shown in FIGS.6 and 7 is a maximum of 3°. The start point and the cut angle areoptimized according to blade and disk geometry to maximize a balancebetween stress reduction on the gas turbine disk 10, stress reduction ofthe gas turbine blade 12, a useful life of the gas turbine blade 12, andmaintaining or improving the aeromechanical behavior of the gas turbineblade. As such, if a dovetail backcut 22 is too large, the backcut willhave a negative effect on the life span of the turbine blade 12. If thedovetail backcut is too small, although the life of the turbine bladewill be maximized, stress concentrations in the interface between theturbine blade and the disk will not be minimized, and the disk would notbenefit from a maximized life span.

The backcut 22 may be planar or as shown in dashed-line in FIG. 6, thebackcut 22′ may alternatively be non-planar. In this context, the cutangle is defined as a starting cut angle. For some turbine classes, thecut angle is pertinent from the start point until the backcut 22, 22′ isdeep enough that the blade loading face of the blade dovetail 16 losescontact with the disk dovetail slot 14. Once contact is lost with thedisk slot 14, any cut of any depth or shape outside the defined envelopewould be acceptable.

As discussed above, where the blade dovetail 16 and disk dovetail slot14 includes a number of tangs 20, a start point and/or cut angle for thedovetail backcut may be determined separately for each of the number oftangs. In a related context, as also referenced above, dovetail backcutsmay be formed in one or both of the pressure side and suction side ofthe turbine blade and/or disk.

Optimization of the start point and cut angle for the dovetail backcutis determined by executing finite element analyses on the blade and diskgeometry. Virtual thermal and structural loads based on engine data areapplied to the blade and disk finite element grids to simulate engineoperating conditions. The no-backcut geometry and a series of varyingbackcut geometries are analyzed using the finite element model. Atransfer function between backcut geometry and blade and disk stressesis inferred from the finite element analyses. The predicted stresses arethen correlated to field data using proprietary materials data in orderto predict blade and disk lives and blade aeromechanical behavior foreach backcut geometry. The optimum backcut geometry and acceptablebackcut geometry range are determined through consideration of both theblade and disk life and the blade aeromechanical behavior.

The datum line W also varies according to blade or disk geometry. Thedatum line W is positioned a fixed distance from a forward face of theblade or disk dovetail along a center line of the dovetail axis. FIGS.21-30 illustrate the datum line W definition for each of the GeneralElectric turbine classes referenced above and for each blade and diskstage. For example, FIG. 21 illustrates the datum line W definition fora stage 1 blade and disk in the first turbine class of a first type (the6FA turbine), where the datum line W is located 1.704 inches from aforward face of the blade and disk dovetail along the center line (datumS) of the dovetail axis. FIG. 22 illustrates the datum line W definitionfor a stage 1 blade and disk in the first turbine class of a second type(the 6FA+e turbine), where the datum line W is located 1.698 inches froma forward face of the blade and disk dovetail along the center line(datum S) of the dovetail axis. FIG. 23 illustrates the datum line Wdefinition for the stage 2 blade and disk in the first turbine class ofthe second type (the 6FA+e turbine), where the datum line W is located1.936 inches from the forward face of the blade and disk dovetail alonga center line (datum S) of the dovetail axis.

FIG. 24 shows the dimension as 2.470 inches for a stage 1 blade and diskin the second turbine class of a first type (the 7FA+e turbine). FIG. 25shows the dimension as 2.817 inches for a stage 2 blade and disk in thesecond turbine class of a second type (the 7FA+e turbine).

FIG. 26 shows the dimension as 2.964 inches for the stage 1 blade anddisk in the third turbine class of a first type (the 9FA+e turbine).FIG. 27 shows the dimension as 3.379 inches for the stage 2 blade anddisk in the third turbine class of the first type (the 9FA+e turbine).FIG. 28 shows the dimension as 2.964 inches for a stage 1 blade in thethird turbine class of a second type (the 9FA turbine).

FIG. 29 shows the dimension as 2.470 inches for a stage 1 blade and diskin the second turbine class of a second type (the 7FA turbine). FIG. 30shows the dimension as 2.817 inches for a stage 2 blade and disk in thesecond turbine class of the second type (the 7FA turbine). The datumline W provides an identifiable reference point for each stage blade anddisk of each turbine class for locating the optimized dovetail backcutstart point.

Details of the optimized start point and cut angle for each turbineclass in each respective blade and disk stage will be described withreference to FIGS. 8-20 and 31-36. As noted, the optimized start pointand cut angle for each dovetail backcut have been determined usingfinite element analyses in order to maximize a balance between stressreduction on the gas turbine disk, stress reduction on the gas turbineblades, a useful life of the gas turbine blades, and maintaining orimproving the aeromechanical behavior of the gas turbine blade. Althoughspecific dimensions will be described, the invention is not necessarilymeant to be limited to such specific dimensions. The maximum dovetailbackcut is measured by the nominal distance to the start point shownfrom the datum line W. Through the finite element analyses, it has beendetermined that a larger dovetail backcut would result in sacrifices tothe acceptable life of the gas turbine blade. In describing the optimaldimensions, separate values may be determined for the number of tangs 20of the blade dovetail 16 and/or the disk dovetail slots 14.

FIGS. 8 and 9 illustrate the values for the stage 1 blade and disk inthe first turbine class of the first type (the 6FA turbine), whichcontains three sets of dovetail tangs here identified by the generalwidth between the tang sets, where the start point of the dovetailbackcut is at least 1.619 inches in an aft direction from the datum lineW for the wide tang, at least 1.552 inches in an aft direction from thedatum line W for the middle tang, and at least 1.419 inches in the aftdirection from the datum line for the narrow tang. The cut angle is amaximum of 3°.

FIGS. 10 and 11 illustrate the values for the stage 1 blade and disk inthe first turbine class of the second type (the 6FA+e turbine), whichcontains three sets of dovetail tangs here identified by the generalwidth between the tang sets, where the start point of the dovetailbackcut is at least 1.549 inches in an aft direction from the datum lineW for the wide tang and the middle tang and at least 1.466 inches in theaft direction from the datum line for the narrow tang. The cut angle isa maximum of 3°. The stage 2 blade and disk in the first turbine classof the second type (the 6FA+e turbine), which contains three sets ofdovetail tangs here identified by the general width between the tangsets is illustrated in FIG. 12. FIG. 12 shows a start point of thedovetail backcut at least 0.923 inches in the aft direction from thedatum line W for the wide tang and at least 1.654 inches in the aftdirection from the datum line W for the middle tang. The cut angle is amaximum of 5°.

FIGS. 13 and 14 illustrate the values for the stage 1 blade and disk inthe second turbine class of the first type (the 7FA+e turbine), whichcontains three sets of dovetail tangs. The start point of the dovetailbackcut is at least 1.945 inches in the aft direction from the datumline, and the cut angle is a maximum of 3°. For the pressure side of thestage 2 blade and disk in the second turbine class of the first type(the 7FA+e turbine), which contains three sets of dovetail tangs hereidentified by the general width between the tang sets, FIG. 15illustrates the start point of the dovetail backcut at least 1.574inches in a forward direction from the datum line W for the wide tang,at least 1.400 inches in the forward direction from the datum line forthe middle tang, and at least 1.226 inches in the forward direction fromthe datum line for the narrow tang. The cut angle is a maximum of 5°.For the suction side of the stage 2 blade and disk in the second turbineclass of the first type (the 7FA+e turbine), which contains three setsof dovetail tangs, as shown in FIG. 16, the start point of the dovetailbackcut is at least 1.725 inches in the aft direction from the datumline. The cut angle is a maximum of 5°. FIGS. 17 and 18 illustrate thestage 1 blade and disk in the third turbine class of the first type (the(9FA+e turbine), which contains three sets of dovetail tangs, where thestart point of the dovetail backcut is at least 1.839 inches in the aftdirection from the datum line W. The cut angle is a maximum of 3°. Thepressure side of the stage 2 blade in the third turbine class of thefirst type (the 9FA+e turbine), which contains three sets of dovetailtangs, is illustrated in FIG. 19. The start point of the dovetailbackcut is at least 1.848 inches in the forward direction from the datumline W. The cut angle is a maximum of 5°. The suction side of the stage2 blade and disk in the third turbine class of the first type (the 9FA+eturbine), which contains three sets of dovetail tangs, is illustrated inFIG. 20. The start point of the dovetail backcut is at least 2.153inches in the aft direction from the datum line W, and the cut angle isa maximum of 5°.

FIGS. 31 and 32 illustrate a stage 1 blade and disk for the second typeof the third turbine class (9FA), according to an exemplary embodimentof the present application. FIG. 31 shows the removal area of thebackcut, which, as illustrated, is located on the suction side aft endof the dovetail. The backcut may be at least approximately 1.539 inchesin the aft direction from the datum line W for each of the threedovetail pressure faces (i.e., tangs). The cut angle may vary betweenabout 0° and 3°. In certain embodiments, as illustrated in FIG. 32, thecut angle may be about 0.7°. Thus, in some embodiments, for example, thebackcut may enter each pressure face at the location described above ata 0.7° angle and then proceed at the 0.7° angle through the remainder ofthe dovetail.

FIGS. 33 and 34 illustrate a stage 1 blade and disk for the second typeof the second turbine class (7FA), according to an exemplary embodimentof the present application. FIG. 33 shows the removal area of thebackcut, which, as illustrated, is located on the suction side aft endof the dovetail. The backcut may be at least approximately 1.645 inchesin the aft direction from the datum line W for each of the threedovetail pressure faces (i.e., tangs). The cut angle may vary betweenabout 0° and 3°. In certain embodiments, as illustrated in FIG. 34, thecut angle may be about 0.7°. Thus, in some embodiments, for example, thebackcut may enter each pressure face at the location described above ata 0.7° angle and then proceed at the 0.7° angle through the remainder ofthe dovetail.

FIGS. 35 and 36 illustrate a stage 2 blade and disk for the second typeof the second turbine class (7FA), according to an exemplary embodimentof the present application. FIG. 35 shows the removal area of thebackcut, which, as illustrated, is located on the suction side aft endof the dovetail. The backcut may begin at least approximately 1.215inches in the aft direction from the datum line W for each of the threedovetail pressure faces (i.e., tangs). The cut angle may vary betweenabout 0° and 3°. In certain embodiments, as illustrated in FIG. 36, thecut angle may be about 2.0°. Thus, in some embodiments, for example, thebackcut may enter each pressure face at the location described above ata 2.0° angle and then proceed at the 2.0° angle through the remainder ofthe dovetail.

It is anticipated that the dovetail backcuts can be formed into a unitduring a normal hot gas path inspection process. With this arrangement,the blade load path should be diverted around the high stress region inthe disk and/or blade stress concentrating features. The relief cutparameters including an optimized start point relative to a datum lineand an optimized cut angle define a dovetail backcut that maximizes abalance between stress reduction in the gas turbine disk, stressreduction in the gas turbine blades, a useful life of the gas turbineblades, and maintaining or improving the aeromechanical behavior of thegas turbine blade. The reduced stress concentrations serve to reducedistress in the gas turbine disk, thereby realizing a significantoverall disk fatigue life benefit.

While the invention has been described in connection with what ispresently considered to be the most practical and preferred embodiments,it is to be understood that the invention is not to be limited to thedisclosed embodiments, but on the contrary, is intended to cover variousmodifications and equivalent arrangements included within the spirit andscope of the appended claims.

1. A method for reducing stress on at least one of a turbine disk and aturbine blade, wherein a plurality of turbine blades are attachable tothe disk, and wherein each of the turbine blades includes a bladedovetail engageable in a correspondingly-shaped dovetail slot in thedisk, the method comprising: (a) determining a start point for adovetail backcut relative to a datum line, the start point defining alength of the dovetail backcut along a dovetail axis; (b) determining acut angle for the dovetail backcut; and (c) removing material from atleast one of the blade dovetail or the disk dovetail slot according tothe start point and the cut angle to form the dovetail backcut; whereinthe start point and the cut angle are optimized according to blade anddisk geometry to maximize a balance between stress reduction on thedisk, stress reduction on the blade, a useful life of the turbineblades, and maintaining or improving the aeromechanical behavior of theturbine blade; wherein the datum line is positioned a fixed distancefrom a forward face of the blade dovetail along a centerline of thedovetail axis, and wherein step (a) is practiced such that the startpoint of the dovetail backcut is at least approximately 1.215 inches inan aft direction from the datum line.
 2. The method of claim 1, whereineach of the turbine blades is configured to operate within a secondstage of a 7FA turbine.
 3. The method according to claim 1, wherein step(b) is practiced such that the cut angle is a maximum of about 3°. 4.The method according to claim 1, wherein step (b) is practiced such thatthe cut angle is about 2°.
 5. The method according to claim 1, whereinthe fixed distance from the forward face of the blade dovetail for thedatum line is approximately 2.817 inches.
 6. The method according toclaim 3, wherein optimizing of the start point and the cut angle ispracticed by executing finite element analyses on the blade and diskgeometry.
 7. The method according to claim 1, wherein step (b) ispracticed by determining multiple cut angles to define the dovetailbackcut with a non-planar surface.
 8. The method according to claim 1,wherein step (c) is practiced by removing material from the bladedovetail.
 9. The method according to claim 1, wherein step (c) ispracticed by removing material from the disk dovetail slot.
 10. Themethod according to claim 1, wherein step (c) is practiced by removingmaterial from the blade dovetail and from the disk dovetail slot. 11.The method according to claim 10, wherein step (c) is further practicedsuch that a resulting angle based on the material removed from the bladedovetail and the disk dovetail slot does not exceed the cut angle.
 12. Aturbine blade comprising an airfoil and a blade dovetail, the bladedovetail being shaped corresponding to a dovetail slot in a turbinedisk; wherein the blade dovetail includes a dovetail backcut sized andpositioned according to blade geometry to maximize a balance betweenstress reduction on the disk, stress reduction on the blade, a usefullife of the turbine blade, and maintaining or improving theaeromechanical behavior of the turbine blade; wherein a start point ofthe dovetail backcut, which defines a length of the dovetail backcutalong a dovetail axis, is determined relative to a datum line positioneda fixed distance from a forward face of the blade dovetail along acenterline of the dovetail axis; and wherein the start point of thedovetail backcut is at least approximately 1.215 inches in an aftdirection from the datum line.
 13. The turbine blade according to claim12, wherein each of the turbine blades is configured to operate within asecond stage of a 7FA turbine.
 14. The turbine blade according to claim12, wherein a cut angle of the dovetail backcut is a maximum of about3°.
 15. The turbine blade according to claim 12, wherein a cut angle ofthe dovetail backcut is about 2°.
 16. A turbine blade according to claim12, wherein the fixed distance from the forward face of the bladedovetail for the datum line is approximately 2.817 inches.
 17. A turbineblade according to claim 12, wherein the dovetail backcut has anon-planar surface.
 18. A turbine blade comprising an airfoil and ablade dovetail, the blade dovetail being shaped corresponding to adovetail slot in a turbine disk; wherein each of the turbine blades isconfigured to operate within a first stage of a 9FA turbine and theblade dovetail includes a dovetail backcut; wherein a start point of thedovetail backcut, which defines a length of the dovetail backcut along adovetail axis, is determined relative to a datum line positioned a fixeddistance from a forward face of the blade dovetail along a centerline ofthe dovetail axis; wherein the start point of the dovetail backcut is atleast 1.215 inches in an aft direction from the datum line; and whereina cut angle of the dovetail backcut is a maximum of about 3°.
 19. Aturbine blade according to claim 17, wherein a cut angle of the dovetailbackcut is about 2°.
 20. A turbine blade according to claim 17, whereinthe fixed distance from the forward face of the blade dovetail for thedatum line is approximately 2.817 inches.